The invention relates to a structural panel component, and particularly a curved structural shell component that is especially for an aircraft fuselage, and that includes at least one skin sheet as well as stiffening elements thereon. The invention further relates to a method of manufacturing such a structural component.
The present day conventional fabrication of large format structural components, and particularly the fuselage shell components for aircraft, typically uses skin sheets having dimensions of approximately 2.5 m by 10 m. The maximum largest possible size of the skin sheets is typically used for fabricating the fuselage shell, in order to minimize the number of the longitudinal and circumferential or transverse joints of the finished fuselage, and thereby reduce the weight and the fabrication effort and expense of the aircraft fuselage. Minimizing the structural weight of the fuselage is a very important factor in the manufacture of aircraft, with regard to the ever present effort to reduce the fuel consumption and therefore improve the economy of operating the aircraft.
Such a weight minimization of the fuselage requires a structure-mechanical optimization of the fuselage shell structure, so that, for example, the skin sheets must have different thicknesses at different areas depending on the respective loads that will prevail at each respective area. This aims to avoid the weight penalty of otherwise having an unnecessary excess thickness at any given area of the fuselage shell that will be subjected to below-average loading. Similarly, areas of the fuselage shell that will be subjected to above-average loading can be provided with a thicker fuselage skin (and/or stronger stringers and frames), without unnecessarily thickening other areas of the fuselage shell. For example, it may be necessary to provide a thickening of the fuselage skin metal sheet in the area of each stringer joint for proper load introduction.
The various different thickness areas of the skin sheet are typically conventionally achieved by riveting or adhesively bonding a doubling or thickening sheet onto the base skin sheet. Alternatively, a reduction of thickness of the fuselage skin can be achieved by mechanical milling or by chemical milling and material removal at these areas. This process of chemical milling and material removal of a skin sheet is carried out by masking the skin sheet, cutting and partially removing portions of the mask at areas that are to be etched, and then applying an etching chemical to remove material from the exposed surface areas, followed by neutralizing, cleaning, etc.
In order to strengthen and support the skin sheet and thereby fabricate the structural shell component, longitudinally extending stringers are typically riveted or adhesively bonded onto the fuselage skin. Then, crosswise or circumferentially extending frames are joined onto the structure being fabricated, by first riveting angle elements, i.e. so-called clips, onto the skin sheet and the stringers. Then the frames are joined onto the clips.
The above described conventional fabrication process for manufacturing a fuselage shell structure is rather complicated, time consuming, and costly. Such a process is described in greater detail in an article by Peter Heider entitled xe2x80x9cLasergerechte Konstruktion und lasergerechte Fertigungsmittel zum Schweissen grossformatiger Aluminium-Strukturbauteilexe2x80x9d (xe2x80x9cLaser Compatible Construction and Laser Compatible Manufacturing Means for Welding Large-Format Aluminum Structural Componentsxe2x80x9d), published by the VDI Verlag Publishers in VDI Fortschritt-Berichte (xe2x80x9cProgress Reportsxe2x80x9d), Series 2: Fertigungstechnik (xe2x80x9cManufacturing Technologyxe2x80x9d), No. 326, Dissertation July/1994, especially at pages 3 to 5.
German Patent DE 198 44 035 discloses a method of fabricating large format structural components, as well as different possible manners of construction of a structural component, which use a laser beam welding process for joining a stiffening structure onto the fuselage skin. Particularly, the stringers running in the longitudinal direction of the aircraft are laser welded onto the skin sheet. From this prior art reference it is also known to embody the frames or frame elements that run in the crosswise or circumferential direction in a weldable manner. Thus, the overall structural component is realized predominantly by welding together numerous individual parts.
FIG. 3 of the present application shows a representative example of a conventionally fabricated aircraft fuselage shell 10, including a skin sheet 11 with stringers 12 arranged and joined thereon. Particularly, the stringers 12 may be riveted or adhesively bonded onto the skin sheet 11. Before carrying out such an adhesive bonding process, all of the individual parts must be subjected to a special pretreatment process, including degreasing, cleaning, pickling or etching, anodizing in a chromic acid solution, and finally being coated with a primer, before the actual adhesive application can take place. Then, a suitable adhesive is applied to the components, which are joined and clamped or held together by appropriate jigs, and then the adhesive is allowed to cure at an elevated temperature and pressure in an autoclave for a sufficient amount of time to ensure adequate and proper bonding of the stringers 12 onto the skin sheet 11. Thereafter, the bonded components must be cleaned, and excess adhesive must be removed. Further, the adhesive joints must be protected against corrosion by applying a bead of an appropriate sealant, and then an additional protective coating to protect against attack by aggressive media.
On the other hand, the above mentioned riveting process for joining the stringers 12 onto the skin sheet 11 is similarly complicated, time-consuming, and costly. Namely, the components must be subjected to a rather complicated preparation process, especially for achieving adequate surface protection. Namely, the skin sheet 11 and the stringers 12 must be anodized, coated with a primer, cleaned with an activator along the joint surfaces, and provided with a sealant along the joint surfaces. Also, a complicated rivet hole boring, cleaning and preparation process must be carried out. Only thereafter can the actual riveting process be carried out. Then the rivets 15 are inserted into the mating holes, and upset or riveted to secure the stringers 12 onto the skin sheet 11.
Thereafter, angle bracket elements, particularly so-called clips 13, are riveted onto the stringers 12 and onto the skin sheet 11. This riveting process also involves the complicated and time consuming steps that were described above. In a further assembly process, the actual crosswise or circumferential frames 14 are joined onto the fuselage shell 10 being fabricated. Particularly, the frame 14 is riveted onto the respective angle bracket elements or clips 13. During this process, the above described complicated, time consuming and costly preparation measures and application of a suitable sealant are necessary at least on certain areas of the fuselage shell or skin.
In view of the above it is an object of the invention to provide a structural component and especially a structural panel or shell component that has a simplified structure and construction, so that it can be fabricated with a simpler, less costly and less time consuming fabrication process, whereby the fabrication time and fabrication costs of the component can be reduced. It is a further object of the invention to provide such a suitable simplified and less costly fabrication process. The invention more particularly aims to avoid complicated riveting and adhesive bonding process and the associated preparatory steps, and to achieve varying dimensions or thicknesses of the structural component at different areas in a simple and economical manner. The invention further aims to avoid or overcome the disadvantages of the prior art, and to achieve additional advantages, as apparent from the present specification.
The above objects have been achieved according to the invention in a structural panel component, especially a shell component for an aircraft fuselage, which comprises at least one skin sheet and a plurality of stiffening elements running in the longitudinal direction and in the crosswise direction on the skin sheet. The term xe2x80x9cstructural panel componentxe2x80x9d encompasses both a flat planar plate or panel component and a curved or contoured shell component. Particularly according to the invention, the longitudinal stiffening elements and the crosswise or transverse stiffening elements are respectively integral with the skin sheet, so that the overall structure including the stiffening elements and the skin sheet is one single monolithic integral piece. According to the invention, this single monolithic integral piece was fabricated by a material chip removing process applied to a suitable semi-finished part. Thus, the finished component has characteristics resulting from having been formed by chip removal machining and particularly milling from a single monolithic solid plate of a starting material. Such characteristics include a fully integral structure, a uniform metal composition and grain pattern and orientation throughout different parts of the component, and a lack of joints, gaps or interruptions, for example.
The above objects have further been achieved according to the invention in a method of fabricating a structural component, wherein a high speed milling process carried out with a large-format or large-area milling machine is used to mill a flat or curved plate-shaped semi-finished part. The milling is carried out according to the proper pattern and requirements to form from the starting plate, a monolithic completely integral structural component including a remaining skin sheet and remaining stiffening elements running integrally along the skin sheet in longitudinal and crosswise or transverse directions.
According to the invention, it is advantageous that the production or fabrication of the structural component using the high speed milling process is optimized from a process point of view, and totally avoids the previously necessary, costly and complicated processes for either adding, thickening or strengthening members or alternatively removing material from the sheet thickness in order to achieve a weight-saving structural component. By using high speed milling to remove material from a semi-finished plate member, it is easy to achieve relatively simply structured, yet very detail-rich structural components, and especially structural shell components. Thereby, various joint locations are also avoided, which minimizes the risks of corrosion and the like.
As a further advantage, the number and variety of separate parts that are needed for fabricating the structural component can be significantly reduced. This in turn significantly simplifies and reduces the effort and complexity of material disposition, storage and warehousing logistics, networking of the construction, tracking of the fabrication or finishing steps, quality control, and the like. Also, when designing new components or the like, relatively few individual parts need to be changed.
According to further detailed or preferred aspects of the invention, the thickness of the remaining skin can be varied in several steps, as necessary, and the dimensions, spacing, etc. of the stiffening elements can be varied as necessary, to appropriately match the resulting local strength of the structural component to the respective loads that will prevail in a given respective area of the structural component. Such variations can be carried out without much effort or expense, simply by providing an appropriate control of the high speed milling process. For example, the process may be controlled by an appropriate computer numerical control (CNC).
Particularly, the stringers can have individual thickness steps, or simply stiffening webs may be provided with any required stepped-down or stepped-up thickness, which is easily achieved with a chip-removing process such as a milling process. A frame element may be joined respectively onto a frame supporting web that is integrally formed with the skin sheet. For example, the frame element may be laser welded onto the frame supporting web. This provides a very strong integral load introduction from the frame into the skin sheet and vice versa. A preferred end use application of the inventive structural component is as a structural shell component of an aircraft fuselage.
The inventive method may be substantially automated, so that the overall fabrication process can be substantially automated. A load optimized component design or structure can easily be achieved, or changed when necessary, by simply providing an appropriate process control, for example through a computer numerical control of the high speed milling process and the laser beam welding process. Such a laser beam welding process will provide a high welding process speed. The combination of the high speed milling and the high speed laser beam welding is thus a very advantageous combination that considerably reduces the fabrication time, effort, and expense.